The present invention relates to gas turbine engines and, more particularly, to aircraft-type high bypass ratio turbine engines having multi-stage compressor and turbine sections.
A typical modern gas turbine aircraft engine, particularly of the high bypass ratio type, includes multi-stage high pressure compressor and turbine sections interconnected by a central compressor shaft or, in some models, a forward shaft. In the latter instance, the forward shaft extends between the webs of the last stage high pressure compressor disk and the fist stage high pressure turbine disk webs. The high pressure turbine section typically includes first and second stage disks in which the second stage disk is attached to the first stage disk by a bolted connection. The interstage volume between the first and second stage disks is enclosed by a shield extending between the out peripheries of the turbine disks. The shield is generally cylindrical in shape and its wall defines an outwardly convex configuration.
The first and second stage disks are isolated by a forward faceplate, attached to the forward face of the first stage disk, and an aft seal attached to the rearward face of the second stage disk web. Typically, cooling air ducted externally from the compressor section is circulated within the volumes defined by the faceplate and aft seal, as well as the interstage volume, in order to cool the disks and the blades they support. The cooling air is conveyed radially outwardly from the turbine section through channels formed in the turbine blades.
In such engines, virtually all of the connections between components are effected through bolting. That is, the forward faceplate is connected to the stage one disk by a circular pattern of bolts extending about the faceplate and disk. The inner periphery of the faceplate is bolted to a disk positioned forwardly of the first stage disk. Similarly, the interstage thermal seal is connected to the turbine disks through bolts in a circular pattern, typically clamping angular blade retaining rims to the opposite faces of the turbine disks as well. In addition, the second stage disk includes a rearwardly-extending cone which is bolted to the aft seal.
A disadvantage with such bolted connections is that they require holes to be formed in the disks which cause stress concentrations and limit the useful lives of the seals and disks. Furthermore, additional disk weight is required to sustain the stresses imposed by the bolt and bolt hole engagement. Accordingly, there is a need for a turbine engine design which minimizes the use of bolted connections between components, yet provides a turbine engine which is relatively easy to assemble and disassemble.
Another disadvantage with such engines is that alignment of the first and second stage disks is difficult to maintain during assembly and operation, which may result in excessive vibrations during operation. Further, in order to convey cooling compressor air to the turbine section, it is necessary to duct the compressor air externally of the turbine and compressor sections. This ducting occupies space in the engine nacelle and adds weight to the engine. Accordingly, there is a need for mounting the first and second stage disks which minimizes alignment problems and further, there is a need for a design which eliminates the need for external ducting of cooling compressor air to the turbine section.